US4916906A - Breach-cooled structure - Google Patents

Breach-cooled structure Download PDF

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Publication number
US4916906A
US4916906A US07/176,482 US17648288A US4916906A US 4916906 A US4916906 A US 4916906A US 17648288 A US17648288 A US 17648288A US 4916906 A US4916906 A US 4916906A
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United States
Prior art keywords
wall
inlets
cooling
breach
cooling fluid
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
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US07/176,482
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English (en)
Inventor
Robert L. Vogt
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General Electric Co
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General Electric Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
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Assigned to GENERAL ELECTRIC COMPANY, A NEW YORK CORP. reassignment GENERAL ELECTRIC COMPANY, A NEW YORK CORP. ASSIGNMENT OF ASSIGNORS INTEREST. Assignors: VOGT, ROBERT L.
Priority to US07/176,482 priority Critical patent/US4916906A/en
Priority to CA000590669A priority patent/CA1327455C/en
Priority to IL89509A priority patent/IL89509A/xx
Priority to DE3908166A priority patent/DE3908166B4/de
Priority to SE8900996A priority patent/SE468060B/sv
Priority to JP1067788A priority patent/JP2783835B2/ja
Priority to FR8903737A priority patent/FR2629134B1/fr
Priority to GB8906805A priority patent/GB2216645B/en
Priority to AU31628/89A priority patent/AU626291B2/en
Priority to IT8919893A priority patent/IT1228872B/it
Priority to US07/489,627 priority patent/US5083422A/en
Publication of US4916906A publication Critical patent/US4916906A/en
Application granted granted Critical
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002—Wall structures
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03044—Impingement cooled combustion chamber walls or subassemblies
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates generally to means and methods for cooling structures in a gas turbine engine, and, more specifically, to an improved cooling method and means usable in a gas turbine engine combustor for reducing exhaust emissions thereof.
  • the prior art includes various means for reducing gas turbine engine exhaust emissions including improved carburetors to more fully mix and atomize fuel and air for obtaining more complete combustion. It is known that unburned hydrocarbons will result where the combustion, or reaction, process occurs at less than about 1500° F., whereas complete burning of hydrocarbons will occur at reaction temperatures greater than about 2000° F., with reaction temperatures therebetween resulting in varying amounts of unburned hydrocarbons.
  • a low temperature boundary layer, or film, of cooling air along the entire inner surface of a combustor liner provides for effective cooling of the liner from the hot combustion gases.
  • the temperature of that boundary layer is about the temperature of the cooling air, which is substantially less than about 1500° F.
  • quenching, or cooling of the fuel/air mixture against that boundary layer will occur during operation. Since combustion of the quenched fuel/air mixture along the cooling air boundary layer will therefore occur at temperatures less than about 1500° F., unburned hydrocarbons and carbon monoxide will be generated.
  • Another object of the present invention is to provide a new and improved combustor effective for reducing exhaust emissions.
  • Another object of the present invention is to provide a new and improved combustor which reduces quenching of the fuel/air mixture.
  • Another object of the present invention is to provide a new and improved cooling means for cooling a wall member.
  • Another object of the present invention is to provide a combustor having new and improved means for cooling a liner thereof to allow for an increase in temperature of combustion gases immediately adjacent a surface of the liner facing the combustion gases.
  • Another object of the present invention is to provide a combustor having new and improved means for cooling a liner portion thereof to allow for the elimination of film cooling of that liner portion to reduce quenching of the fuel/air mixture for reducing exhaust emissions.
  • the present invention comprises a method and apparatus for improving cooling of a wall in a gas turbine engine.
  • the method and apparatus provide breach cooling of an imperforate wall, such as a combustor liner, for improving the cooling thereof.
  • the breach cooling is effected by structure for channeling a cooling fluid as a jet toward an outer surface of the imperforate wall, with the jet having sufficient momentum to breach a boundary layer of the cooling fluid which forms over the wall outer surface to allow the jet to contact the wall outer surface for more effective cooling thereof.
  • the breach-cooled wall is an upstream portion of the gas turbine engine combustor, and the inner surface of the combustor liner facing the combustion gases is characterized by not having a film-cooling boundary layer of air to reduce quenching of the combustion gases for reducing exhaust emissions.
  • FIG. 1 is a sectional view of a gas turbine engine combustor in accordance with one embodiment of the present invention.
  • FIG. 2 is a plan view of a portion of the first panel illustrated in FIG. 1 taken along line 2--2.
  • FIG. 3 is a sectional view of an enlarged portion of a first liner panel of the combustor illustrated in FIG. 1 in accordance with one embodiment of the present invention taken along line 3--3 of FIG. 2.
  • FIG. 4 is an end, partly sectional, view of the first panel illustrated in FIG. 3 taken along line 4--4.
  • FIG. 5 is a partly sectional schematic view of one of the inlets illustrated in FIG. 3 illustrating a flow-turning skirt in accordance with one embodiment of the present invention.
  • FIG. 6 is a partly sectional perspective view of a hollow gas turbine engine blade including a breach-cooled concave wall in accordance with another embodiment of the present invention.
  • FIG. 1 Illustrated in FIG. 1 is a sectional view of an annular gas turbine engine combustor 10 disposed concentrically about an engine centerline or axial axis 12. Disposed upstream of the combustor 10 is a conventional compressor (not shown) which provides compressed air, or cooling fluid, 14 to the combustor 10.
  • the combustor 10 includes an annular outer liner 16 having an outer surface 18 facing and spaced from an annular casing 20 to define an annular first passage 22 therebetween for receiving a portion of the cooling fluid 14.
  • the combustor 10 also includes an annular inner liner 24 having an outer surface 26 spaced from and facing an inner casing 28 to define an annular second passage 30 therebetween for receiving a portion of the cooling fluid 14.
  • the inner liner 24 is spaced from the outer liner 16 to define a combustion zone 32 therebetween.
  • the outer liner 16 and the inner liner 24 include oppositely facing inner surfaces 34 and 36, respectively, which bound the combustion zone 32.
  • the inner and outer liners 24, 16 are formed of a conventional metallic material such as commercially available Hastelloy X.
  • the combustor 10 also includes an annular dome 38 fixedly joined to upstream ends 40 and 42 of the outer liner 16 and inner liner 24, respectively.
  • the combustor 10 also includes a plurality of circumferentially-spaced carburetors 44 disposed in the dome 38, with each including a conventional low pressure, airblast fuel injector 46 and a conventional counterrotating swirler assembly 48.
  • Fuel 50 is discharged from the injector 46 into the swirler 48 and mixed with the air 14 to create a fuel/air mixture 52, which is discharged from an outlet 54 of each carburetor 44 into the combustion zone 32 for burning, or reacting, therein and generating combustion, or exhaust, gases 56.
  • the exhaust gases 56 flow through the combustor 10 generally along a longitudinal axis 58 thereof and are discharged from a combustor outlet 60, from which they are then channeled to a turbine (not shown).
  • Each of the outer and inner liners 16 and 24 includes a plurality of serially-joined liner panels including a first liner panel 62 extending downstream from the dome 38 and followed, in turn, by a second liner panel 64 and a third liner panel 66. Disposed at the intersections of the first, second and third liner panels 62, 64 and 66 and at the downstream end of the third liner panel 66 are conventional film-cooling nuggets 68, each of which generally includes an aft-facing U-shaped pocket having an aperture in the base thereof for discharging the air 14.
  • the cooling nuggets 68 are effective for channeling the cooling air 14 from the apertures as a boundary layer film 70 downstream along the inner surfaces 34 and 36 of the liners for providing film cooling thereof.
  • the inventor has discovered that instead of using a relatively large number of relatively small conventional impingement cooling apertures, i.e. increasing the density of such apertures, for increasing the heat transfer coefficient of the cooling fluid, the use of a reduced number of relatively large cooling apertures can result in a substantial increase in heat transfer coefficient for cooling the liner.
  • the inventor has discovered a new method of and structure for cooling a wall for obtaining an increase in heat transfer therefrom. This is accomplished by a method of "breach” cooling an imperforate wall wherein a cooling fluid is channeled as a jet toward the outer surface of the wall with the jet having sufficient momentum to breach, i.e. penetrate, the boundary layer of cooling fluid which is formed over the wall outer surface, with the jet contacting the wall outer surface.
  • the cooling air jet does not simply contact and merge with the boundary layer itself, but breaches that boundary layer to directly contact the wall outer surface.
  • the first liner panel 62 of the outer liner 16 Illustrated in more particularly in FIG. 3 is the first liner panel 62 of the outer liner 16, it being understood that a substantially identical, but inverted, panel 62 is also found in the inner liner 24.
  • the first liner panel 62 is a breach-cooled structure including an imperforate inner, or first, wall 72 extending from the dome 38, at the upstream liner end 40, to the second liner panel 64.
  • the inner wall 72 includes an inner surface 74 which faces the combustion zone 32, and an opposing outer, or first, surface 76.
  • the first liner panel 62 further includes an aft wall 78 extending outwardly from the inner wall 72 at an aft end thereof.
  • An outer, or second, wall 80 is spaced from the inner wall 72 and is suitably fixedly joined at an aft end thereof to the aft wall 78 and at an upstream end to the inner wall 72 itself to define an enclosed plenum 82 therebetween.
  • the outer wall 80 includes a plurality of inlet apertures, or simply inlets, 84 which face the outer surface 76.
  • the aft wall 78 includes a plurality of circumferentially-spaced outlet apertures, or simply outlets, 86 for discharging the cooling fluid 14 from the plenum 82.
  • the first liner panel 62 further includes means for channeling the cooling fluid 14 from each of the inlets 84 as a jet 88 (i.e. jets 88a, 88b, 88c) across the plenum 82 to breach a boundary layer of air 90 which forms along the outer surface 76 to make direct contact with the outer surface 76.
  • the channeling means includes, inter alia, the inlets and outlets 84 and 86 being sized and positioned so that the momentum (mass times velocity) of the cooling fluid jet 88 flowable through the inlets 84 is sufficient to breach the boundary layer 90 to contact the outer surface 76.
  • annular flow-turning skirt 92 is preferably attached to or formed integral with each inlet 84.
  • the skirt 92 preferably extends below the outer wall 80 a minimum of about two wall thicknesses thereof for effectively turning the cooling air 14 and directing the jets 88.
  • the skirts 92 were effective for directing the jets 88 toward the inner wall 72 at an angle A measured therebetween of about 75°.
  • FIG. 3 illustrates that during operation of the combustor 10, the cooling fluid 14 at the inlets 84 is at a static pressure P 1 , which represents compressor discharge pressure.
  • the static pressure of the cooling fluid 14 in the plenum 82 at the outlet 86 has a value P 2 .
  • the static pressure of the cooling fluid 14 in the cooling nuggets 68 downstream of the outlet 86 has a value P 3 , which represents turbine inlet static pressure.
  • the total pressure drop, P 1 minus P 3 designated P 13 , is a given value and is controlled by the engine cycle.
  • a first pressure drop P 1 minus P 2 is made predeterminedly larger than a second pressure drop P 2 minus P 3 , designated P 23 , to assist in ensuring breach cooling of the inner wall 72. Since the total static pressure drop P 13 is a given, the inlets 84 and the outlets 86 may be sized so that the first pressure drop P 12 has a maximum value and the second pressure drop P 23 has a minimum value to ensure that all of the jets 88 breach the boundary layer 90.
  • the inlets 84 include a first, upstream row of circumferentially-spaced first inlet apertures 84a, a second row of circumferentially-spaced second inlet apertures 84b spaced longitudinally downstream of the first inlets 84a and followed, in turn, by a third row of circumferentially-spaced third, aft inlet apertures 84c.
  • Each of the inlets 84a, 84b and 84c has a flow area A 1 , A 2 , A 3 , respectively, and is spaced perpendicularly from the outer surface 76 a distance H 1 , H 2 , and H 3 , respectively.
  • a first jet 88a of cooling fluid 14 is generated by the first inlet 84a to contact a first, upstream portion 76a of the outer surface 76 and forms the boundary layer 90, which has a momentum and a thickness T 1 , and is flowable downstream over a second portion 76b of the outer surface 76, which faces the second inlet 84b.
  • the flow area A 2 of the second inlet 84b along with the momentum of a second jet 88b generated by the second inlet 84b are predeterminedly selected relative to the thickness T 1 and momentum of the boundary layer 90 to ensure that the second jet 88b breaches the boundary layer 90 to contact the outer surface second portion 76b.
  • the flow area A 3 of the third inlet 84c and the momentum of a third jet 88c generated by the third inlet 84c are predeterminedly selected relative to the momentum and thickness T 2 of the boundary layer 90, which flows over the outer surface third portion 76c to ensure that the third jet 88c breaches the boundary layer 90 to contact the outer surface 76.
  • the second inlets 84b are circumferentially spaced between the first inlets 84a and the third inlets 84c as illustrated in FIG. 3.
  • the inlets 84 had a uniform 0.115 inch diameter and were spaced from each other 0.28 inch in the longitudinal direction between rows and 0.77 inch in the circumferential direction.
  • the heights H 1 , H 2 and H 3 were 0.25 inch, 0.30 inch, and 0.35 inch, respectively.
  • the total static pressure P 13 was about 14 psi, and the first pressure drop P 12 was about 11 psi and the second pressure drop P 23 was about 3 psi.
  • the additional parameter C is a constant
  • m is a variable
  • Re is the conventionally known Reynolds number
  • Pr is the conventionally known Prandtl number which is substantially constant for the cooling fluid 14 at a given temperature
  • H is the height H 1 , H 2 , or H 3 of the inlets 84.
  • the parameters C and m are conventionally determined from empirical data generated for particular applications and are a function of the thickness and momentum of the boundary layer 90.
  • ⁇ is the density of the cooling fluid 14
  • V is the velocity of the jet 88
  • ⁇ is the absolute coefficient of viscosity of the cooling fluid 14.
  • the additional parameter W is the mass flow rate of the cooling fluid 14 through each row of the inlets 84
  • a e is the conventionally known effective area of each row of the inlets 84 (e.g. A 1 , A 2 , A 3 ).
  • the additional parameter A et is the sum of the effective area of all the inlet rows and W t is the total mass flow rate.
  • the additional parameter A eo is the overall effective flow area of the inlets 84 and outlets 86.
  • the additional parameter A ed is the sum of the effective areas of the outlets 86.
  • a heat transfer coefficient h of about 600 btu/hr-ft 2 -°F. is typical of conventional impingement cooling.
  • a heat transfer coefficient h of about 900 btu/hr-ft 2 -°F. was obtained for the preferred embodiment of breach cooling illustrated in FIG. 3. This represents a substantial 50% improvement.
  • the substantial benefit of breach cooling is further emphasized in comparison to conventional film cooling of a back surface of an imperforate liner resulting in a typical heat transfer coefficient of about only 300 btu/hr-ft 2 -°F.
  • h avg is the average heat transfer coefficient due to a particular row
  • K and c are constants
  • N is the total number of rows of the inlets 84.
  • K would be equal to the average heat transfer coefficient for that row.
  • Equations 8 and 9 represent the effect of the increasing thickness of the boundary layer 90.
  • the breach cooling effect will be lost, resulting more closely in a conventional impingement-cooled structure.
  • three rows of the inlets 84 were found to be effective for obtaining breach cooling.
  • the inner wall 72 is spaced from and overlaps the second liner panel 64 to define the cooling nugget 68, which includes a downstream end of the inner wall 72, a portion of the aft wall 78 including the outlets 86, and an upstream portion of the second liner panel 64.
  • the cooling fluid 14 is discharged from the plenum 82 through the outlets 86 and is caused by the cooling nuggets 68 to flow as the boundary layer film 70 of cooling air along the inner surface 34 of the outer liner 16 in a conventionally known manner.
  • the breach-cooled structure i.e. first liner panel 62 as substantially illustrated in FIG. 3, was tested and was found to be effective for obtaining acceptably low temperatures of inner wall 72 by dissipating heat from the inner wall 72, which had a heat flux rate of about 450,000 btu/hr-ft 2 , since no film cooling on the combustion side of the wall was utilized.
  • This is a substantial increase in heat flux when compared to the heat flux of about 150,000 btu/hr-ft 2 , which is typically found in a conventional liner provided with primarily film cooling of the surface of the liner facing the combustion zone.
  • the combustor 10 can be operated without film cooling of the inner surface of the first panel 62, thus increasing the temperature which the fuel/air mixture 52 experiences along the inner wall 72 and resulting in reduced exhaust emissions.
  • the inner surface was effectively operated at about 1500° F. without the need for film cooling thereof.
  • a preferred embodiment of the invention further includes a thermal barrier coating 94 over the entire inner surface 74 of the inner wall 72.
  • the coating 94 is conventional, and 8% yittrium-zirconia was used in the preferred embodiment with a thickness of about 10 mils.
  • the heat flux through the inner wall 72 was reduced to about 300,000 btu/hr-ft 2 .
  • the surface of the coating 94 facing the combustion zone 32 experienced a temperature of about 1900° F. in the preferred embodiment tested, and, at this temperature, the unburned hydrocarbons and carbon monoxide were substantially reduced so that the combustor 10 was able to meet the more stringent recently issued FAA and ICAO emissions requirements.
  • the breach-cooled first liner panel 62 is utilized solely as the liner immediately downstream from the dome 38 for several reasons. Although the breach-cooled liner panel 62 has a high heat transfer coefficient, a relatively large differential in temperature exists from the inner surface 74 to the outer surface 76. Large temperature gradients are generally undesirable because they may shorten the useful life of a structure. Furthermore, inasmuch as combustion occurs primarily in the combustion zone 32 in the vicinity immediately downstream of the dome 38 and generally within the first liner panel 62, breach cooling of the first liner panel 62 only was found effective for adequately reducing the exhaust emissions to meet the FAA and ICAO requirements. Accordingly, the second and third liner panels 64 and 66 remain basically unchanged from the original and conventional film-cooling design.
  • first liner panel 62 is operated without film cooling of the inner surface for reducing exhaust emissions, it was found that conventional cooling, including impingement cooling, was not adequate to effectively cool an imperforate liner wall. It was also found that only breach cooling in accordance with the invention was effective for adequately cooling the imperforate inner wall 72.
  • FIG. 6 illustrated in FIG. 6 is a hollow gas turbine engine rotor blade 96 through which the cooling fluid 14 flows.
  • the blade 96 includes an imperforate concave outer, or first, wall 98 having an inner, first surface 100, and an outer, second surface 102 over which the combustion gases flow.
  • An insert, or second wall, 104 is disposed in the blade 96 and is spaced from the inner surface 100 to define a plenum 106.
  • the insert 104 includes three axially-spaced rows, each including a plurality of radially-spaced inlets 108.
  • the blade 96 includes a plurality of radially-spaced trailing edge outlets 110 which discharge cooling fluid from the plenum 106.
  • the insert 104, concave wall 98, inlets 108 and outlets 110 are generally similar to the breach-cooling structure illustrated in FIG. 2.
  • the inlets 108 and the outlets 110 are predeterminedly sized and configured to provide breach cooling of the concave side 98 of the blade 96.
  • any liner portion or an entire combustor liner may be breach-cooled, depending on particular design requirements.
  • one embodiment tested suggests that three rows of inlets 84 provide optimum cooling, the actual limit on the number of rows is limited only by the ability to achieve the improved heat transfer coefficient from the breach cooling disclosed herein.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US07/176,482 1988-03-25 1988-03-25 Breach-cooled structure Expired - Lifetime US4916906A (en)

Priority Applications (11)

Application Number Priority Date Filing Date Title
US07/176,482 US4916906A (en) 1988-03-25 1988-03-25 Breach-cooled structure
CA000590669A CA1327455C (en) 1988-03-25 1989-02-09 Breach-cooled structure
IL89509A IL89509A (en) 1988-03-25 1989-03-07 Breach-cooled structure
DE3908166A DE3908166B4 (de) 1988-03-25 1989-03-13 Prallgekühltes Gebilde
SE8900996A SE468060B (sv) 1988-03-25 1989-03-21 Anordning foer att kyla en konstruktion
FR8903737A FR2629134B1 (fr) 1988-03-25 1989-03-22 Procede de refroidissement par rupture et structure ainsi refroidie
JP1067788A JP2783835B2 (ja) 1988-03-25 1989-03-22 ブリーチ冷却構造及びガスタービンエンジン燃焼器
GB8906805A GB2216645B (en) 1988-03-25 1989-03-23 Cooling of wall members of structures
AU31628/89A AU626291B2 (en) 1988-03-25 1989-03-23 Breach-cooled structure
IT8919893A IT1228872B (it) 1988-03-25 1989-03-24 Struttura raffreddata a breccia.
US07/489,627 US5083422A (en) 1988-03-25 1990-03-07 Method of breach cooling

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US07/176,482 US4916906A (en) 1988-03-25 1988-03-25 Breach-cooled structure

Related Child Applications (1)

Application Number Title Priority Date Filing Date
US07/489,627 Division US5083422A (en) 1988-03-25 1990-03-07 Method of breach cooling

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US4916906A true US4916906A (en) 1990-04-17

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US07/176,482 Expired - Lifetime US4916906A (en) 1988-03-25 1988-03-25 Breach-cooled structure

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US (1) US4916906A (de)
JP (1) JP2783835B2 (de)
AU (1) AU626291B2 (de)
CA (1) CA1327455C (de)
DE (1) DE3908166B4 (de)
FR (1) FR2629134B1 (de)
GB (1) GB2216645B (de)
IL (1) IL89509A (de)
IT (1) IT1228872B (de)
SE (1) SE468060B (de)

Cited By (33)

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US5083422A (en) * 1988-03-25 1992-01-28 General Electric Company Method of breach cooling
US5123248A (en) * 1990-03-28 1992-06-23 General Electric Company Low emissions combustor
US5224822A (en) * 1991-05-13 1993-07-06 General Electric Company Integral turbine nozzle support and discourager seal
US5363654A (en) * 1993-05-10 1994-11-15 General Electric Company Recuperative impingement cooling of jet engine components
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US5388412A (en) * 1992-11-27 1995-02-14 Asea Brown Boveri Ltd. Gas turbine combustion chamber with impingement cooling tubes
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US5749229A (en) * 1995-10-13 1998-05-12 General Electric Company Thermal spreading combustor liner
US6389815B1 (en) 2000-09-08 2002-05-21 General Electric Company Fuel nozzle assembly for reduced exhaust emissions
US6749396B2 (en) 2002-06-17 2004-06-15 General Electric Company Failsafe film cooled wall
US20040250548A1 (en) * 2003-06-11 2004-12-16 Howell Stephen John Floating liner combustor
US20060137324A1 (en) * 2004-12-29 2006-06-29 United Technologies Corporation Inner plenum dual wall liner
US20060283189A1 (en) * 2005-06-15 2006-12-21 General Electric Company Axial flow sleeve for a turbine combustor and methods of introducing flow sleeve air
US20070283700A1 (en) * 2006-06-09 2007-12-13 Miklos Gerendas Gas-turbine combustion chamber wall for a lean-burning gas-turbine combustion chamber
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US20080271458A1 (en) * 2007-03-01 2008-11-06 Srinath Varadarajan Ekkad Zero-Cross-Flow Impingement Via An Array of Differing Length, Extended Ports
US20090282833A1 (en) * 2008-05-13 2009-11-19 General Electric Company Method and apparatus for cooling and dilution tuning a gas turbine combustor liner and transition piece interface
US20100037622A1 (en) * 2008-08-18 2010-02-18 General Electric Company Contoured Impingement Sleeve Holes
US20100251723A1 (en) * 2007-01-09 2010-10-07 Wei Chen Thimble, sleeve, and method for cooling a combustor assembly
US20110107769A1 (en) * 2009-11-09 2011-05-12 General Electric Company Impingement insert for a turbomachine injector
US20110214429A1 (en) * 2010-03-02 2011-09-08 General Electric Company Angled vanes in combustor flow sleeve
US20120304659A1 (en) * 2011-03-15 2012-12-06 General Electric Company Impingement sleeve and methods for designing and forming impingement sleeve
US20130081401A1 (en) * 2011-09-30 2013-04-04 Solar Turbines Incorporated Impingement cooling of combustor liners
US9010125B2 (en) 2013-08-01 2015-04-21 Siemens Energy, Inc. Regeneratively cooled transition duct with transversely buffered impingement nozzles
US20150121885A1 (en) * 2013-11-05 2015-05-07 Mitsubishi Hitachi Power Systems, Ltd. Gas Turbine Combustor
US9411016B2 (en) 2010-12-17 2016-08-09 Ge Aviation Systems Limited Testing of a transient voltage protection device
US20160265772A1 (en) * 2013-11-04 2016-09-15 United Technologies Corporation Turbine engine combustor heat shield with multi-height rails
US11371701B1 (en) 2021-02-03 2022-06-28 General Electric Company Combustor for a gas turbine engine
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SE8900996L (sv) 1989-09-26
DE3908166A1 (de) 1989-10-05
DE3908166B4 (de) 2007-11-08
JP2783835B2 (ja) 1998-08-06
GB2216645A (en) 1989-10-11
SE8900996D0 (sv) 1989-03-21
JPH01301929A (ja) 1989-12-06
AU626291B2 (en) 1992-07-30
SE468060B (sv) 1992-10-26
GB2216645B (en) 1992-09-02
FR2629134B1 (fr) 1994-07-08
IT8919893A0 (it) 1989-03-24
AU3162889A (en) 1989-09-28
IL89509A (en) 1992-06-21
CA1327455C (en) 1994-03-08
IT1228872B (it) 1991-07-05
FR2629134A1 (fr) 1989-09-29
IL89509A0 (en) 1989-09-10
GB8906805D0 (en) 1989-05-10

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